Gas turbine engine component cooling cavity with vortex promoting features

ABSTRACT

A component according to an exemplary aspect of the present disclosure includes, among other things, a body, a wall extending inside of the body and a plurality of vortex promoting features arranged in a helical pattern along the wall.

CROSS-REFERENCED TO RELATED APPLICATION

This application is a divisional of U.S. application Ser. No.15/102,335, which was filed on Jun. 7, 2016, which is a National PhaseApplication of International Application No. PCT/US14/65858, which wasfiled on Nov. 17, 2014, which claims priority to U.S. ProvisionalApplication No. 61/919,286, which was filed Dec. 20, 2013.

BACKGROUND

This disclosure relates to a gas turbine engine, and more particularlyto a gas turbine engine component that includes a cooling cavity. Thecooling cavity may be provided with a plurality of vortex promotingfeatures that induce a vortexing flow of a cooling fluid through thecavity.

Gas turbine engines typically include a compressor section, a combustorsection, and a turbine section. During operation, air is pressurized inthe compressor section and is mixed with fuel and burned in thecombustor section to generate hot combustion gases. The hot combustiongases are communicated through the turbine section, which extractsenergy from the hot combustion gases to power the compressor section andother gas turbine engine loads.

Due to their exposure to hot combustion gases, many gas turbine enginecomponents employ internal cooling schemes that channel a dedicatedcooling fluid for cooling the component. Thermal energy is transferredfrom the component to the cooling fluid as it circulates through thecooling scheme to convectively cool the component. Some cooling schemesmay additionally rely on film cooling holes that return a portion of thecooling fluid to the gas path as a layer of film that protects thecomponent against the relatively harsh environment of the gas path.

SUMMARY

A component according to an exemplary aspect of the present disclosureincludes, among other things, a body, a wall extending inside of thebody and a plurality of vortex promoting features arranged in a helicalpattern along the wall.

In a further non-limiting embodiment of the foregoing component, thecomponent is one of a blade, a vane, a blade outer air seal (BOAS), acombustor liner and a turbine exhaust case liner.

In a further non-limiting embodiment of either of the foregoingcomponents, the wall circumscribes a cavity formed through the body.

In a further non-limiting embodiment of any of the foregoing components,the helical pattern extends along a direction of flow through thecavity.

In a further non-limiting embodiment of any of the foregoing components,the cavity includes a bowed shape.

In a further non-limiting embodiment of any of the foregoing components,the cavity includes a serpentine shape.

In a further non-limiting embodiment of any of the foregoing components,the cavity includes a helical shape.

In a further non-limiting embodiment of any of the foregoing components,the helical pattern is singularly helical.

In a further non-limiting embodiment of any of the foregoing components,the helical pattern is dually helical.

In a further non-limiting embodiment of any of the foregoing components,the helical pattern includes a first helix that extends in parallel witha second helix.

In a further non-limiting embodiment of any of the foregoing components,the helical pattern is multiply helical.

In a further non-limiting embodiment of any of the foregoing components,the plurality of vortex promoting features include at least one ofpedestals, fins, ribs and hemispherical protrusions.

A gas turbine engine according to another exemplary aspect of thepresent disclosure includes, among other things, a component thatdefines a cavity configured to channel a cooling fluid to cool thecomponent. The component includes a wall that at least partiallycircumscribes the cavity and a plurality of vortex promoting featuresarranged in a helical pattern along the wall and configured to induce avortexing flow of the cooling fluid through the cavity.

In a further non-limiting embodiment of the foregoing gas turbineengine, the helical pattern extends along a direction of flow throughthe cavity.

In a further non-limiting embodiment of either of the foregoing gasturbine engines, the cavity includes one of a bowed, serpentine,helical, conical or contoured shape.

A method according to another exemplary aspect of the present disclosureincludes, among other things, manufacturing a component to include aplurality of vortex promoting features arranged in a helical patternalong a wall of the component.

In a further non-limiting embodiment of the foregoing method, the methodof manufacturing includes additively manufacturing a casting articlelayer-by-layer using an additive manufacturing process and casting thecomponent to include the plurality of vortex promoting features usingthe casting article.

In a further non-limiting embodiment of either of the foregoing methods,the method of manufacturing includes building the component to includethe plurality of vortex promoting features layer-by-layer using anadditive manufacturing process.

In a further non-limiting embodiment of any of the foregoing methods,the component is made of a ceramic material.

In a further non-limiting embodiment of any of the foregoing methods,the component is made of a refractory metal.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

The various features and advantages of this disclosure will becomeapparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 illustrates a schematic, cross-sectional view of a gas turbineengine.

FIGS. 2A and 2B illustrate a first embodiment of a gas turbine enginecomponent.

FIGS. 3A and 3B illustrate a second embodiment of a gas turbine enginecomponent.

FIG. 4 illustrates an exemplary cooling scheme that may be employed by agas turbine engine component.

FIG. 5 illustrates a cooling scheme according to a second embodiment ofthis disclosure.

FIG. 6 illustrates a cooling scheme according to yet another embodimentof this disclosure.

FIG. 7 illustrates vortex promoting features of a cooling scheme.

FIG. 8 illustrates a second embodiment of vortex promoting features.

FIG. 9 illustrates another embodiment of vortex promoting features.

FIG. 10 yet another embodiment of vortex promoting features.

FIG. 11 illustrates a cavity of a cooling scheme.

FIG. 12 illustrates another cavity of a cooling scheme.

FIG. 13 illustrates yet another cavity of a cooling scheme.

FIG. 14 illustrates a casting article that can be used to manufacture agas turbine engine component that includes a cooling scheme with vortexpromoting features.

DETAILED DESCRIPTION

This disclosure is directed to a gas turbine engine component thatincludes one or more cooling cavities that employ vortex promotingfeatures for inducing a vortexing fluid flow within the cavity. Thevortex promoting features may be arranged in a helical pattern along thedirection of flow through the cavity. Rotation of the cooling fluidalong the helical path promotes a relatively high convective heattransfer coefficient, thereby removing additional amounts of heat fromthe component. The gas turbine engine component may be additivelymanufactured to include the vortex promoting features, or may bemanufactured using an additively manufactured casting article that isconfigured to create the helical pattern inside the component during acasting process. These and other features are discussed herein.

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmenter section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram ° R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

Each of the compressor section 24 and the turbine section 28 may includealternating rows of rotor assemblies and vane assemblies (shownschematically). For example, the rotor assemblies can carry a pluralityof rotating blades 25, while each vane assembly can carry a plurality ofvanes 27 that extend into the core flow path C. The blades 25 create orextract energy in the form of pressure from the core airflow that iscommunicated along the core flow path C. The vanes 27 direct the coreairflow to the blades 25 to either add or extract energy.

Various components of the gas turbine engine 20, including but notlimited to the airfoils of the blades 25 and the vanes 27, may besubjected to repetitive thermal cycling under widely rangingtemperatures and pressures. The hardware of the turbine section 28 isparticularly subjected to relatively extreme operating conditions.Therefore, some components may require internal cooling schemes forcooling the parts during engine operation. Exemplary embodiments of suchcooling schemes are discussed in greater detail below.

FIGS. 2A and 2B illustrate a component 60A having an internal coolingscheme 65 for cooling the component 60A. The component 60A may include abody 62 that defines both an external and internal shape of thecomponent 60A. The body 62 extends between a leading edge LE and atrailing edge TE and may include an airfoil section 64, a platform 66and a root 68. The airfoil section 64 extends outwardly in a firstdirection from the platform 66, and the root 68 extends from theplatform 66 in an opposed, second direction away from the airfoilsection 64.

The exemplary cooling scheme 65 may include one or more cavities 70 thatare disposed through the body 62. The cavities 70 are shownschematically via dashed lines in FIG. 2A, and it is to be understoodthat the cooling scheme 65 is not limited to the number of cavities 70depicted by these figures. Although not shown, the cavity 70 may be fedwith a cooling fluid, such as relatively cool bleed airflow sourced fromthe compressor section 24, for convectively cooling the component 60A.

In one embodiment, the cavities 70 are main cooling cavities of thecooling scheme 65. The cavities 70 may extend radially, axially andcircumferentially inside of the body 62.

In the embodiment illustrated by FIGS. 2A and 2B, the component 60A is ablade, such as a turbine blade. It is to be understood; however, thatthe cooling schemes described herein are not limited for use in turbineblades and can be also employed within vanes, blade outer air seals(BOAS), combustor liners, casing structures, turbine exhaust caseliners, or any other gas turbine engine component that might benefitfrom dedicated internal cooling, including some compressor components.

For example, a second non-limiting embodiment of a component 60B thatmay employ a cooling scheme 65 is illustrated by FIGS. 3A and 3B. Inthis embodiment, the component 60B is a vane. The component 60B mayinclude a body 71 that includes an airfoil section 73 that extendsbetween an inner platform 75 and an outer platform 77.

As discussed in greater detail below, the cooling schemes 65 describedabove may include a plurality of vortex promoting features that arearranged in a helical pattern to promote a vortexing flow of the coolingfluid communicated through the cavities 70. The vortexing flow of thecooling fluid through the cavities 70 promotes a higher convective heattransfer coefficient, thereby removing a greater amount of heat from thecomponent and potentially reducing the need for film cooling.

FIG. 4 illustrates a portion of a cooling scheme 65 that includes acavity 70 disposed inside of a gas turbine engine component 85. The gasturbine engine component 85 could be a blade or vane (similar to thoseshown in FIGS. 2A, 2B and 3A, 3B), or any other gas turbine enginecomponent that may benefit from dedicated internal cooling.

In one embodiment, the gas turbine engine component 85 includes a wall72 extending inside of a body 87. The wall 72 circumscribes the cavity70 of the cooling scheme 65. For ease of illustration, only a singlecavity 70 has been illustrated. In other words, the cooling scheme 65could include many additional features, including cavities,microcircuits, etc., that are not shown in this particular embodiment.

A plurality of vortex promoting features 74 may be formed or otherwisedisposed on the wall 72. In one embodiment, the vortex promotingfeatures 74 are arranged in a helical pattern 76 along the wall 72. Thehelical pattern 76 may extend in a direction of flow 78 of the coolingfluid communicated within the cavity 70. In other words, the vortexpromoting features 74 are implemented circumferentially in the cavity70, varying in location along the direction of flow 78 within the cavity70, to form the helical pattern 76. The helical pattern 76 may stretchover a portion of the cavity 70, or over the entire length of the cavity70.

In one embodiment, the helical pattern 76 may be singularly helical (seeFIG. 4). In another embodiment, the helical pattern 76 may be duallyhelical (see FIG. 5). A dually helical pattern includes two spaced aparthelixes H1 and H2 of vortex promoting features 74 that extend generallyparallel to one another. In yet another embodiment, the helical pattern76 may be made up of a plurality of helixes H1 to H_(n) of vortexpromoting features 74 (see FIG. 6). In other words, the helical pattern76 may be made up of any number helixes that extend relative to oneanother in composition. The helical pattern 76 may have a constant orvariable pitch and circumference, and can be arranged to extend ineither a clockwise or counterclockwise direction.

The vortex promoting features 74 may also embody a variety of sizes,shapes and configurations. For example, by way of non-limitingembodiments, the vortex promoting features 74 may include pedestals (seeFIG. 7), hemispherical protrusions (see FIG. 8), fins (see FIG. 9), ribs(see FIG. 10), or any other features. The views depicted in FIGS. 7-10are similar to a view through section A-A of FIGS. 2B and 3B, forexample.

The cavity 70 of the cooling scheme 65 could also embody any of avariety of shapes and configurations. For example, the cavity 70depicted in FIG. 4 includes a straight shape. However, the cavity 70could alternatively include a bowed shape (see FIG. 11), a serpentineshape (see FIG. 12), or a helical shape (see FIG. 13). Other threedimensional shapes, including but not limited to contoured shapes andconical shapes, are also contemplated as within the scope of thisdisclosure. In combination with the helical pattern 76 of vortexpromoting features 74, the shape of the cavity 70 may further contributeto the creation of an internal vortex of fluid flow within the cavity70, thereby improving heat transfer.

FIG. 14 illustrates an exemplary casting article 100 that can beadditively manufactured and subsequently used in a casting operation tocast a gas turbine engine component (e.g., gas turbine engine components60A, 60B, 85 or any other gas turbine engine component) that includesvortex promoting features that are disposed in a helical pattern, suchas illustrated in FIGS. 4-10.

The exemplary casting article 100 includes a body 102 having a pluralityof indents 104 that extend into the body 102. In one embodiment, theindents 104 are arranged in a helical pattern in order to form vortexpromoting features disposed in a helical pattern along a wall of a castgas turbine engine component.

In one embodiment, the casting article 100 is a casting core. However, ashell, gating, or other casting articles may also be additivelymanufactured and used to create a gas turbine engine component havingvortex promoting features that are disposed in a helical pattern.

In one non-limiting additive manufacturing process, the casting article100 is built layer-by-layer by delivering a powdered material, such as aceramic material or refractory metal material, to a build platform. Alayer of the powdered material is then melted at locations where thegeometry of the casting article 100 is to exist. A laser, electron beammelting device or any other melting device may be used to melt thelayers of powdered material. The build platform may then be moved and asecond layer of powered material may be added and melted to prepare asecond layer of the casting article. This layer-by-layer process may berepeated until the entire casting article 100 has been additively built.

In one embodiment, the layers of the casting article 100 may be joinedto one another with reference to CAD data that defines a cross-sectionof a desired geometry of the casting article 100. Once additivelymanufactured, the casting article 100 may be used in a casting processto manufacture a gas turbine engine component having vortex promotingfeatures arranged in a helical pattern. Alternatively, in anotherembodiment, the entirety of the gas turbine engine component may beadditively manufactured to include vortex promoting features arranged ina helical pattern.

Although the different non-limiting embodiments are illustrated ashaving specific components, the embodiments of this disclosure are notlimited to those particular combinations. It is possible to use some ofthe components or features from any of the non-limiting embodiments incombination with features or components from any of the othernon-limiting embodiments.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed and illustrated in these exemplary embodiments,other arrangements could also benefit from the teachings of thisdisclosure.

The foregoing description shall be interpreted as illustrative and notin any limiting sense. A worker of ordinary skill in the art wouldunderstand that certain modifications could come within the scope ofthis disclosure. For these reasons, the following claims should bestudied to determine the true scope and content of this disclosure.

1. A method, comprising: additively manufacturing a casting article layer-by-layer using an additive manufacturing process; and casting a component to include a plurality of vortex promoting features arranged in a helical pattern along a wall of the component; using the casting article.
 2. The method as recited in claim 1, wherein the component is made of a ceramic material.
 3. The method as recited in claim 1, wherein the component is made of a refractory metal.
 4. The method as recited in claim 1, wherein the casting article includes a body having a plurality of indents extending into the body, and the indents are arranged in a helical pattern to form the plurality of vortex promoting features.
 5. The method as recited in claim 1, wherein the step of manufacturing includes building the component to include the plurality of vortex promoting features layer-by-layer using an additive manufacturing process.
 6. The method as recited in claim 5, wherein the casting article is a casting core.
 7. The method as recited in claim 1, wherein the casting article is a casting core.
 8. The method as recited in claim 1, wherein the additive manufacturing step includes building the casting article layer by layer by delivering a powdered metal to a build platform, and melting a layer of powdered material.
 9. The method as recited in claim 8, comprising: moving the build platform; adding a second layer of powdered material; and melting the second layer.
 10. The method as recited in claim 8, wherein the melting is performed by a laser.
 11. The method as recited in claim 8, wherein the melting is performed by an electron beam melting device.
 12. The method as recited in claim 8, wherein the powdered material is a ceramic material.
 13. The method as recited in claim 8, wherein the powdered material is a refractory metal material. 